Methods and apparatuses for aerial interception of aerial threats

ABSTRACT

Embodiments include active protection systems and methods for an aerial platform. An onboard system includes radar modules, detects aerial vehicles within a threat range of the aerial platform, and determines if any of the aerial vehicles are an aerial threat. The onboard system also determines an intercept vector to the aerial threat, communicates the intercept vector to an eject vehicle, and causes the eject vehicle to be ejected from the aerial platform to intercept the aerial threat. The eject vehicle includes alignment thrusters to rotate a longitudinal axis of the eject vehicle to substantially align with the intercept vector, a rocket motor to accelerate the eject vehicle along an intercept vector, divert thrusters to divert the eject vehicle in a direction substantially perpendicular to the intercept vector, and attitude control thrusters to make adjustments to the attitude of the eject vehicle.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.16/584,378, filed Sep. 26, 2019, which will issue as U.S. Pat. No.11,313,650 on Apr. 26, 2022, which is a continuation of U.S. patentapplication Ser. No. 15/411,324, filed Jan. 20, 2017, now U.S. Pat. No.10,436,554, issued Oct. 8, 2019, which is a continuation of U.S. patentapplication Ser. No. 13/839,637, filed Mar. 15, 2013, now U.S. Pat. No.9,551,552, issued Jan. 24, 2017, which is a continuation-in-part of U.S.patent application Ser. No. 13/455,831, filed Apr. 25, 2012, now U.S.Pat. No. 9,170,070, issued Oct. 27, 2015, which claims priority to U.S.Provisional Patent Application Ser. No. 61/606,010, filed Mar. 2, 2012,the disclosure of each of which is hereby incorporated herein in itsentirety by this reference. This application is also related to U.S.patent application Ser. No. 13/839,176, filed Mar. 15, 2013, now U.S.Pat. No. 9,501,055, issued Nov. 22, 2016, and titled “Methods andApparatuses for Engagement Management of Aerial Threats,” the disclosureof each of which is hereby incorporated herein in its entirety by thisreference.

TECHNICAL FIELD

Embodiments of the present disclosure relate generally to methods andapparatuses for engagement management relative to a threat and, moreparticularly, to aerial interception of aerial threats.

BACKGROUND

Rocket-propelled grenades (RPGs) and other human carried projectilessuch as man-portable air-defense systems (MANPADS or MPADS) andshoulder-launched surface-to-air missiles (SAMs) represent seriousthreats to mobile land and aerial platforms. Even inexperienced RPGoperators can engage a stationary target effectively from 150-300meters, while experienced users could kill a target at up to 500 meters,and moving targets at 300 meters. One known way of protecting a platformagainst RPGs is often referred to as active protection and generallycauses explosion or discharge of a warhead on the RPG at a safe distanceaway from the threatened platform. Other known protection approachesagainst RPGs and short-range missiles are more passive and generallyemploy fitting the platform to be protected with armor (e.g., reactivearmor, hybrid armor or slat armor).

Active protection systems (APS) have been proposed for ground vehiclesfor defense against RPGs and other rocket fired devices with a goodsuccess rate for quite some time. However, these systems are proposed toprotect vehicles that are: 1) armored, 2) can carry heavy loads, and 3)have plenty of available space for incorporation of large criticalsystems. Currently these systems can weigh anywhere between 300 to 3000lbs. and can protect the vehicle when intercepting incoming threats asclose as 5 to 10 ft.

There is a need in the art for engagement management systems that canwork in cooperation with intercept vehicles to engage and destroy aerialthreats. There is also a need for such systems to be portable andlightweight enough for carrying on aerial and other mobile platformsthat may have significant weight and size constraints, or on which anactive protection system may be easily installed. There is also a needfor such systems to coordinate with multiple engagements of aerialthreats, intercept vehicles, and other nearby engagement managementsystems.

BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1A and 1B illustrate a helicopter as an aerial platform that maybe under attack from an aerial threat and coverage areas that may beemployed to sense when such a threat is present;

FIGS. 2A and 2B illustrate a conventional dispenser in which an ejectvehicle (EV) according to one or more embodiments of the presentdisclosure may be placed;

FIG. 3 illustrates systems that may be present on a helicopter and thatmay intercommunicate according to one or more embodiments of the presentdisclosure;

FIG. 4 illustrates an exploded view of an eject vehicle showing variouselements of the EV according to one or more embodiments of the presentdisclosure;

FIGS. 5A-5C illustrate the eject vehicle of FIG. 4 as it may beconfigured during various stages of an intercept mission according toone or more embodiments of the present disclosure;

FIGS. 6A-6C illustrate various propulsion and thruster elements that maybe included with one or more embodiments of the present disclosure;

FIG. 7 illustrates various electrical and communication connections thatmay be present on an EV while it is disposed on the mobile platformprior to launch;

FIG. 8 is a block diagram illustrating elements that may be present onthe eject vehicle according to one or more embodiments of the presentdisclosure;

FIG. 9A is a block diagram illustrating elements that may be present onthe aerial platform according to one or more embodiments of the presentdisclosure;

FIG. 9B is a perspective view of a radar module that may be present onthe aerial platform according to one or more embodiments of the presentdisclosure;

FIGS. 10A and 10B are diagrams illustrating radar scanning beams duringan acquisition mode and a tracking mode, respectively;

FIG. 11 is a spectrum diagram illustrating possible Doppler spectrumregions where various aerial vehicles may be detected;

FIG. 12 is a simplified flow diagram illustrating some of the processesinvolved in one or more embodiments of the present disclosure;

FIG. 13 illustrates an example flight path for the eject vehicle and anaerial threat during an intercept process;

FIG. 14 illustrates two aerial vehicles flying in a formation andvarious radar sectors that may be covered by the aerial vehicles;

FIG. 15 is a simplified side view of a kill vehicle illustrating theprinciple forces of interest on the kill vehicle;

FIGS. 16A and 16B illustrate a nose thruster module;

FIG. 17 is a simplified block diagram of an attitude control loop;

FIG. 18 is a simplified flowchart of a joint adaptive polarization androll angle estimator (JAPRAE) algorithm;

FIG. 19 illustrates the JAPRAE algorithm in the form of a conventionalservo loop;

FIG. 20 illustrates a non-limiting example of a set of simulationresults of the JAPRAE algorithm of FIGS. 18 and 19 ;

FIG. 21 illustrates a set of position results of an extended Kalmanfilter (EKF) fusion algorithm simulation;

FIG. 22 illustrates a set of velocity results of the EKF fusionalgorithm simulation of FIG. 21 ;

FIG. 23 illustrates a set of attitude results of the EKF fusionalgorithm simulation of FIG. 21 ;

FIG. 24 illustrates an attitude control loop;

FIG. 25 illustrates a controller input-output topology;

FIG. 26 illustrates an example of a search procedure that may be used tofind a thruster to fire; and

FIG. 27 illustrates a mass and center of gravity (CG) location updatealgorithm.

DETAILED DESCRIPTION

In the following description, reference is made to the accompanyingdrawings in which is shown, by way of illustration, specific embodimentsof the present disclosure. The embodiments are intended to describeaspects of the disclosure in sufficient detail to enable those skilledin the art to practice the invention. Other embodiments may be utilizedand changes may be made without departing from the scope of thedisclosure. The following detailed description is not to be taken in alimiting sense, and the scope of the present invention is defined onlyby the appended claims.

Furthermore, specific implementations shown and described are onlyexamples and should not be construed as the only way to implement orpartition the present disclosure into functional elements unlessspecified otherwise herein. It will be readily apparent to one ofordinary skill in the art that the various embodiments of the presentdisclosure may be practiced by numerous other partitioning solutions.

In the following description, elements, circuits, and functions may beshown in block diagram form in order not to obscure the presentdisclosure in unnecessary detail. Additionally, block definitions andpartitioning of logic between various blocks is exemplary of a specificimplementation. It will be readily apparent to one of ordinary skill inthe art that the present disclosure may be practiced by numerous otherpartitioning solutions. Those of ordinary skill in the art wouldunderstand that information and signals may be represented using any ofa variety of different technologies and techniques. For example, data,instructions, commands, information, signals, bits, symbols, and chipsthat may be referenced throughout the description may be represented byvoltages, currents, electromagnetic waves, magnetic fields or particles,optical fields or particles, or any combination thereof. Some drawingsmay illustrate signals as a single signal for clarity of presentationand description. It will be understood by a person of ordinary skill inthe art that the signal may represent a bus of signals, wherein the busmay have a variety of bit widths and the present disclosure may beimplemented on any number of data signals including a single datasignal.

The various illustrative logical blocks, modules, and circuits describedin connection with the embodiments disclosed herein may be implementedor performed with a general-purpose processor, a special-purposeprocessor, a digital signal processor (DSP), an application specificintegrated circuit (ASIC), a field programmable gate array (FPGA) orother programmable logic device, discrete gate or transistor logic,discrete hardware components, or any combination thereof designed toperform the functions described herein. A general-purpose processor maybe a microprocessor, but in the alternative, the processor may be anyconventional processor, controller, microcontroller, or state machine. Ageneral-purpose processor may be considered a special-purpose processorwhile the general-purpose processor is configured to executeinstructions (e.g., software code) stored on a computer-readable medium.A processor may also be implemented as a combination of computingdevices, such as a combination of a DSP and a microprocessor, aplurality of microprocessors, one or more microprocessors in conjunctionwith a DSP core, or any other such configuration.

In addition, it is noted that the embodiments may be described in termsof a process that may be depicted as a flowchart, a flow diagram, astructure diagram, or a block diagram. Although a process may describeoperational acts as a sequential process, many of these acts can beperformed in another sequence, in parallel, or substantiallyconcurrently. In addition, the order of the acts may be rearranged.

Elements described herein may include multiple instances of the sameelement. These elements may be generically indicated by a numericaldesignator (e.g., 110) and specifically indicated by the numericalindicator followed by an alphabetic designator (e.g., 110A) or a numericindicator preceded by a “dash” (e.g., 110-1). For ease of following thedescription, for the most part, element number indicators begin with thenumber of the drawing on which the elements are introduced or most fullydiscussed. For example, where feasible, elements in FIG. 3 aredesignated with a format of 3xx, where 3 indicates FIG. 3 and xxdesignates the unique element.

It should be understood that any reference to an element herein using adesignation such as “first,” “second,” and so forth does not limit thequantity or order of those elements, unless such limitation isexplicitly stated. Rather, these designations may be used herein as aconvenient method of distinguishing between two or more elements orinstances of an element. Thus, a reference to first and second elementsdoes not mean that only two elements may be employed or that the firstelement must precede the second element in some manner. In addition,unless stated otherwise, a set of elements may comprise one or moreelements.

Embodiments of the present disclosure include apparatuses and methodsfor providing protection for mobile platforms, such as, for example, ahelicopter, from an aerial threat. Some embodiments of the presentdisclosure may include methods and apparatuses that are portable andlightweight enough for carrying on aerial platforms that may havesignificant weight and size constraints. Some embodiments of the presentdisclosure may include methods and apparatuses that can be incorporatedinto existing systems already installed on aerial platforms.

FIGS. 1A and 1B illustrate a helicopter as an aerial platform 100 thatmay be under attack from an aerial threat 120 and coverage areas 140that may be employed to sense when such a threat is present within anintercept range (may also be referred to herein as a threat range) ofembodiments of the present disclosure. As shown in FIG. 1A, the aerialthreat 120 may be shot by an attacker 110 toward the aerial platform100.

As used herein, “aerial threat” or “threat” are used interchangeably torefer to any threat directed toward a mobile platform, includingprojectiles, rockets, and missiles that may be shoulder launched orlaunched from other platforms. As non-limiting examples, such aerialthreats include rocket-propelled grenades (RPGs), man-portableair-defense systems (MANPADS or MPADS), shoulder-launched surface-to-airmissiles (SAMs), tube-launched, optically tracked, wire-guided missiles(TOWs), and other aerial weapons, having a trajectory and ordnance suchthat they may cause damage to the mobile platform.

The term “aerial platform” includes, but is not limited to, platformssuch as helicopters, unmanned airborne vehicles (UAVs), remotely pilotedvehicles (RPVs), light aircraft, hovering platforms, and low speedtraveling platforms. The protection systems and methods of the presentdisclosure are particularly useful for protecting aerial platformsagainst many kinds of aerial threats.

While embodiments of the present disclosure may be particularly suitablefor use on aerial platforms 100 due to the small size and weight, theymay also be used in other types of mobile platforms like ground-basedmobile platforms such as, for example, tanks, armored personnelcarriers, personnel carriers (e.g., Humvee and Stryker vehicles) andother mobile platforms capable of bearing embodiments of the presentdisclosure. Moreover, embodiments of the present disclosure may be usedfor relatively stationary ground-based personnel protection wherein amobile platform may not be involved. Accordingly, embodiments of thedisclosure are not limited to aerial applications.

FIG. 1B illustrates coverage areas 140 in which one or more embodimentsof the present disclosure may detect an incoming aerial threat 120 andperform active countermeasures using one or more embodiments of thepresent invention to remove the aerial threat 120 before it can damagethe aerial platform 100. Some embodiments of the present disclosure maybe configured such that they can be disposed in previously existingcountermeasures dispenser systems (CMDS).

FIGS. 2A and 2B illustrate a dispenser 200 configured as a conventionalCMDS (e.g., an AN/ALE-47) in which an eject vehicle 400 (EV) accordingto one or more embodiments of the present disclosure may be placed.AN/ALE-47 dispensers are conventionally used to dispense passivecountermeasures, such as, for example, radar-reflecting chaff, infraredcountermeasures to confuse heat-seeking missile guidance, and disposableradar transmitters. With some embodiments of the present disclosure,eject vehicles 400 may also be placed in the AN/ALE-47 and ejectedtherefrom under control of the AN/ALE-47 and other electronics on theaerial platform 100 (FIGS. 1A and 1B). The eject vehicle 400 may beconfigured as a substantially cylindrical vehicle to be placed in atubular dispenser 210 and ejection may be controlled from control wiring220 connected to the dispenser 200. Moreover, the dispenser 200 may beconfigured to hold both the passive countermeasures for which it wasoriginally designed, as well as one or more eject vehicles 400 accordingto embodiments of the present disclosure.

While some embodiments of the eject vehicle 400 may be configured to bedisposed in an AN/ALE-47, other types of dispensers 200 or other typesof carriers for the eject vehicle 400 may also be used. Moreover, thetubular dispenser 210 is illustrated with a circular cross section.However, other cross sections may be used, such as, for example, square,hexagonal, or octagonal.

FIG. 3 illustrates systems that may be present on a helicopter frame 300and that may intercommunicate according to one or more embodiments ofthe present disclosure. The helicopter frame 300 and systems describedare used as specific examples to assist in giving details aboutembodiments of the present disclosure. In the specific example of FIG. 3, an AAR-47 missile approach warning system (MAWS) warns of threatmissile approaches by detecting radiation associated with the missile.In the specific example, four MAWSs (320A, 320B, 320C, and 320D) aredisposed near four corners of the helicopter frame 300. A centralprocessor 360 may be used to control and coordinate the four MAWSs(320A, 320B, 320C, and 320D).

Two AN/ALE-47 dispensers (200A and 200B) are positioned on outboardsides of the helicopter frame 300, each of which may contain one or moreeject vehicles 400. As shown in FIG. 3 , there are four eject vehicles400 on each side labeled EV1 through EV4 on one side and labeled EV5-EV8on the other side. The AN/ALE-47 dispensers (200A and 200B) are eachcontrolled by an AN/ALE-47 sequencer (350A and 350B), which are, inturn, controlled by the central processor 360.

According to one or more embodiments of the present disclosure fourradar modules (900A, 900B, 900C, and 900D) are included to augment andconnect with the AAR-47s and communicate with the eject vehicles 400.These radar modules 900 (see FIG. 9A) are configured to detect and trackrelatively small incoming aerial threats (e.g., an RPG) as well as theoutgoing eject vehicles 400. Moreover, the radar modules 900 can sendwireless communications (340A, 340B, 340C, and 340D) to the ejectvehicles 400 both before and after they are ejected from the dispensers(200A and 200B). The radar modules 900, and eject vehicles 400 may eachinclude unique identifiers, such as, for example, a media access control(MAC) address. The radar modules 900 may also be configured to detect,track, and communicate with other friendly platforms such as, forexample, other helicopters flying in formation with the helicopter.Thus, all helicopters within communication range can communicate andshare radar and control information to form a broad coverage area,similar to cellular telephone base station coverage. Moreover, and asexplained more fully below, the helicopters may communicate to definedifferent sector coverage areas such that one helicopter does not launchan eject vehicle 400 into a sector that may damage or interfere withanother helicopter.

The control processors, such as the central processor 360, the MAWSs320, the radar modules 900, the sequencers 350, and the dispensers 200may be configured to form an ad hoc network and include the ejectvehicles 400.

The specific example of FIG. 3 is shown to illustrate how radar modules(900A-900D) and eject vehicles (EV1-EV8) of the present disclosure canbe incorporated with existing systems on helicopter platforms withlittle change. Of course, other systems may be employed with embodimentsof the present disclosure. As a non-limiting example, one radar module900A may be positioned on one side of the helicopter frame 300 andanother radar module 900C may be positioned on another side of thehelicopter frame. In such a case, the radar modules 900 would beconfigured to provide hemispherical coverage areas. These radar modules900 may be controlled by, communicate with, or a combination thereof, adifferent central processor 360 configured specifically for embodimentsof the present disclosure. Moreover, the eject vehicles 400 may bedisposed in different carriers or different dispensers from theAN/ALE-47 dispensers (200A and 200B) shown in FIG. 3 .

When embodiments of the present disclosure are used as illustrated inFIG. 3 , they provide an ultra-lightweight active protection system forhelicopter platforms that may increase the survivability against RPGattacks to better than 90% for RPGs fired from ranges as close as about100 meters away.

In order to satisfy the helicopter platform constraints, embodiments ofthe present disclosure address many significant technology areas:

1) For helicopter applications, size, weight, and power should beconsidered. Every pound of added airframe equipment will reduce capacityto carry personnel or cargo, and the space for adding equipment to theairframe may be at a premium. At least some embodiments of the presentdisclosure are configured to be less than about 50 pounds and occupyabout 5.5 inches×5.5 inches surface area at each of the four corners ofa helicopter exterior shell and with minimal impact to existing wiringkits.

2) Helicopters generally do not carry armor and thus, the intercept ofan incoming threat (e.g., an RPG) must occur at a range that is safe tothe un-armored helicopter airframe. Using an RPG-7 as an example, toachieve a survival probability of about 99% from the blast alone, theintercept should occur at distances beyond 30 meters from the helicoptershell. This requirement significantly influences the system responsetime, when considering that an RPG fired at a 100-meter distance mayimpact the helicopter in less than about 600 milliseconds.

3) A third concern is fratricide and collateral damage to friendlyforces that may be amplified by the helicopter platform deployingkinetic countermeasures in a position above ground and potentially nextto a wingman helicopter or in the vicinity of civilians, friendlytroops, or a combination thereof. Some embodiments of the presentdisclosure are configured to work in combination with embodiments onother helicopters when the helicopters are flying in formationrelatively close to each other.

4) Some embodiments of the present disclosure can geo-locate theattacker 110 (FIG. 1A) after a few radar track frames are processed.

5) Embodiments of the present disclosure can engage multiple threats ata time. In other words, multiple incoming aerial threats 120 can bedetected and tracked and multiple outgoing eject vehicles 400 can betracked. In addition, to increase a probability of destroying anincoming aerial threat 120, multiple eject vehicles 400 may be launched,directed toward, and detonated proximate the same aerial threat 120.

6) Finally, eject vehicles 400 can be launched and guided to the pointof attack with the same or different warheads and detonated above thethreat point of origin.

To address these technology areas, some embodiments of the presentdisclosure include an active kinetic countermeasure projectile (i.e.,the eject vehicle 400 of FIG. 2B), including an ejection mechanism withan impulse charge that can fit in, and can be launched by, the AN/ALE-47chaff/flare dispenser 200. Some embodiments of the present disclosureinclude the radar module 900 covering a 90-degree sector or more (i.e.,with a 90-degree sector, each helicopter platform would use four radarmodules 900).

When referring to the radar module 900 herein (e.g., as shown in FIG. 3), it should be understood that in some embodiments the radar module 900may perform the operations described herein in combination with otherelectronics and processors on the aerial platform 100. As such, theradar modules 900 may be used to: 1) search, acquire, and track incomingaerial threats 120, 2) launch the active kinetic countermeasure (i.e.,eject vehicle 400), 3) track the outgoing eject vehicle 400 with respectto the incoming aerial threat 120, 4) point and guide the eject vehicle400 toward the incoming aerial threat 120, 5) command detonate the ejectvehicle 400, and 6) geo-locate the attacker 110, all in less than aboutone second. In one configuration, at least two AN/ALE-47 dispensers 200would be used in conjunction with the four radar modules 900 such thateach dispenser 200 provides hemispherical coverage.

The radar modules 900 may be configured as pulse Doppler radar modules900 to scan the azimuth plane and the elevation plane using twoorthogonal fan beams and may be configured to cover a 90-degree sectorin about 20 milliseconds. Upon detecting an incoming aerial threat 120,the associated radar module 900 may then direct the launch and guidanceof an eject vehicle 400 from an AN/ALE-47 dispenser 200 that covers thatsector. The eject vehicle 400 may be command guided to the target by theradar module 900 and command detonated. The radar modules 900 may beconfigured as an addition to the existing AN/AAR-47 system and may useits existing interface for launching of the eject vehicle 400.

Some of the embodiments of the present disclosure may be configured todeploy an eject vehicle 400 that fits in a standard dispenser 200 butcould be stabilized and pointed towards the threat after launch, in lessthan about 50 milliseconds, in the rotor downwash of a helicopter, andwhen ejected in the fixed direction dictated by the dispenser 200. Theradar modules 900 may then guide the eject vehicle 400 to accuratelyintercept the aerial threat 120 within about 330 milliseconds and thusreduce the requirement of carrying a large warhead.

FIG. 4 illustrates an exploded view of an eject vehicle 400 showingvarious elements of the eject vehicle 400 according to one or moreembodiments of the present disclosure. Reference may also be made toFIGS. 1A-3 in describing features and operations of the eject vehicle400. The eject vehicle 400 is a lightweight guided projectile that, insome embodiments, may be designed to be launched from chaff/flaredispensers. The eject vehicle 400 may intercept and destroy incomingaerial threats 120 at ranges sufficient to prevent damage to the hostaerial platform 100. The eject vehicle 400 may be packaged in acartridge containing an impulse charge and interface electronicsdesigned to fit the AN/ALE-47 dispenser magazine.

The eject vehicle 400 includes an ejection piston 780 configured totransmit the energy of an impulse cartridge 750 (described below inconnection with FIG. 7 ) to the eject vehicle 400 and launch the ejectvehicle 400 away from the aerial platform 100 to a distance safe enoughfor the eject vehicle 400 to begin performing alignment and interceptionmaneuvers.

A rocket motor 420 may be used to propel the eject vehicle 400 towardthe aerial threat 120 after the eject vehicle 400 has been rotated suchthat a longitudinal axis of the eject vehicle 400 is pointed in thegeneral direction of the aerial threat 120. A first set of folding fins482 may be attached to the rocket motor 420 and configured to deployonce the eject vehicle 400 has exited the dispenser 200. The foldingfins 482 are small and configured to provide stability to the ejectvehicle 400 during its flight path rather than as control surfaces fordirecting the fight path.

An airframe shell 430 may be configured to contain a warhead 440, adivert thruster module 610, a nose thruster module 620 (may also bereferred to herein as an alignment thruster module 620), an electronicsmodule 450, and a battery 452. An airframe nose 490 may be configured toattach to the airframe shell 430 to protect the electronics module 450and provide a somewhat aerodynamic nose for the eject vehicle 400.

A safe and arm module 460 may be included within the airframe shell 430and configured to safely arm the warhead 440 when the eject vehicle 400is a safe distance away from the aerial platform 100.

FIGS. 5A-5C illustrate the eject vehicle 400 of FIG. 4 as it may beconfigured during various stages of an intercept mission according toone or more embodiments of the present disclosure. Stage 1, in FIG. 5A,illustrates the eject vehicle 400 in a cartridge 710 (FIG. 7 ) andincludes the ejection piston 780, the rocket motor 420, the airframeshell 430, and the airframe nose 490.

Stage 2, in FIG. 5B, illustrates the eject vehicle 400 after it has beendispensed and shows the rocket motor 420, the airframe shell 430, andthe airframe nose 490. FIG. 5B also illustrates the folding fins 482deployed near the end of the rocket motor 420 and wireless communicationantennas 890 deployed near the airframe nose 490.

Stage 3, in FIG. 5C illustrates the eject vehicle 400 after the rocketmotor 420 has burned and been detached from the airframe shell 430. Atthis stage, the eject vehicle 400 may be referred to as a terminalvehicle and includes the airframe nose 490, the wireless communicationantennas 890, and the airframe shell 430. Still within the airframeshell 430 are the warhead 440, the divert thruster module 610, thealignment thruster module 620, the electronics module 450, the battery452, and the safe and arm module 460. After the rocket motor 420 isdetached, a second set of folding fins 484 are deployed from theairframe shell 430 to stabilize the eject vehicle 400 during theremainder of the flight to intercept the aerial threat 120. This secondset of folding fins 484 are used to replace the first set of foldingfins 482 that were attached to the rocket motor 420, which has beendetached from the airframe shell 430 during stage 3.

In addition, after the rocket motor 420 is detached, one or more cornerreflectors 470 are exposed. The corner reflector 470 may be configuredwith sharp angles to enhance radar detection of the eject vehicle 400 bya radar module 900 on the aerial platform 100. For example, the cornerreflector 470 may be configured as an interior angle of a small cubeshape, which will enhance radar detection.

Returning to FIG. 4 , the alignment thruster module 620 is offset from acenter of mass of the eject vehicle 400 such that an initial pitchmaneuver can be performed to align the longitudinal axis of the ejectvehicle 400 along an intercept vector pointed toward the aerial threat120. This alignment maneuver is performed prior to the burn of therocket motor 420.

The divert thruster module 610 is positioned substantially near a centerof mass of the terminal vehicle and is used to laterally divert theterminal vehicle from its current flight path to make minor correctionsto the flight path in order to more accurately intercept the aerialthreat 120. The terminal vehicle may be referred to herein as the ejectvehicle 400 and it should be understood what is being referred to basedon the context of the discussion.

The warhead 440 may be command detonated when the radar module 900 onthe aerial platform 100 determines that the eject vehicle 400 hasreached the closest point of approach (nominally about 15 cm). The useof thrusters, provide the fast reaction times that may be needed tointercept the aerial threat 120 at a nominal distance of about 50 meterswhen the aerial threat 120 is launched from a range of about 100 meters.

FIGS. 6A-6C illustrate various propulsion and thruster elements that maybe included with one or more embodiments of the present disclosure. FIG.6A illustrates a nose thruster module 620 with four nose thrusters 622(two are hidden) arranged around a periphery of the nose thruster module620. These nose thrusters 622 (also referred to herein as alignmentthrusters 622) are positioned to generate a perpendicular force on theeject vehicle 400 relative to the longitudinal axis and are offset fromthe center of mass of the eject vehicle 400 so that an initial pitchmaneuver can be performed to rotate and align the longitudinal axis ofthe eject vehicle 400 along an intercept vector pointed toward theaerial threat 120. In this embodiment, the four nose thrusters 622 areorthogonally arranged giving two opportunities to adjust the pitch ofthe eject vehicle 400 in each direction. Of course, other embodimentsmay include fewer or more alignment thrusters 622.

FIG. 6B illustrates a divert thruster module 610 with eight divertthrusters 612 (five are hidden) arranged around a periphery of thedivert thruster module 610. These divert thrusters 612 are positioned togenerate a perpendicular force on the eject vehicle 400 relative to thelongitudinal axis and are positioned near the center of mass of theeject vehicle 400 so that the divert thrusters 612 will move the ejectvehicle 400 laterally to a slightly different travel path, whilesubstantially maintaining the same pitch. Thus, the divert thrusters 612can modify the flight path of the eject vehicle 400 to correct for minorerrors in the initial pitch maneuvers pointing directly toward theaerial threat 120. In this embodiment, eight divert thrusters 612 areused giving eight opportunities to adjust the flight path of the ejectvehicle 400 during its flight toward the aerial threat 120. Of course,other embodiments may include fewer or more divert thrusters 612.

FIG. 6C illustrates a thruster 650 configured to expel a gas through anozzle 652 to create a lateral force. The thruster 650 may be controlledfrom a thrust signal 654, which may be connected to the electronicsmodule 450 of the eject vehicle 400. The thruster 650 is one example ofa type of thruster that may be used for both the divert thrusters 612and the alignment thrusters 622.

FIG. 7 illustrates various electrical and communication connections thatmay be present on the eject vehicle 400 while it is disposed on theaerial platform 100 (FIGS. 1A and 2B) prior to launch. A cartridge 710includes a cartridge flange 720 such that the cartridge 710 may besecurely placed in a dispenser 200 (FIG. 2A). An end cap 790 may bepositioned over the cartridge 710 to hold the eject vehicle 400 withinthe cartridge 710. An impulse cartridge 750 is positioned near the baseof the cartridge flange 720 and is configured to fire in response to afire command signal 755 from the radar module 900 (FIG. 3 ) or otherelectronics on the aerial platform 100. An ejection piston 780 ispositioned between the impulse cartridge 750 and the eject vehicle 400and is configured to transmit the energy of the firing impulse cartridge750 to the eject vehicle 400 and propel the eject vehicle 400 out of thedispenser 200 and safely away from the aerial platform 100.

A power signal 740 and a ground signal 730 may run along or through thecartridge to an antenna spring contact 745 and a ground spring contact735, respectively. The ground spring contact 735 is configured toflexibly couple with a ground patch 738 on the eject vehicle 400 toprovide a ground for the eject vehicle 400 electronics while the ejectvehicle 400 is in the cartridge 710. The antenna spring contact 745 isconfigured to flexibly couple with the antenna 890 on the eject vehicle400 and a power signal on the eject vehicle 400 to provide power anddirect communication for the eject vehicle 400 electronics while theeject vehicle 400 is in the cartridge 710. The cartridge 710 may includea cartridge antenna 760 that may be coupled to the antenna 890 of theeject vehicle 400 by the antenna spring contact 745. Thus, the ejectvehicle 400 may communicate wirelessly 795 with electronics onboard theaerial platform 100 through the antenna 890 on the eject vehicle 400 orthrough the cartridge antenna 760.

FIG. 8 is a block diagram illustrating elements that may be present onthe eject vehicle 400 according to one or more embodiments of thepresent disclosure. A microcontroller 810 may be coupled to a memory820, which is configured to hold instructions for execution by themicrocontroller 810 and data related to command and control of the ejectvehicle 400. The microcontroller 810 may be any suitablemicrocontroller, microprocessor, or custom logic configured to directlyexecute, or execute responsive to software instructions, processesrelated to operation of the eject vehicle 400. The memory 820 may be anysuitable combination of volatile and non-volatile memory configured tohold data and computing instructions related to operation of the ejectvehicle 400.

One or more antennas 890 may be configured to provide a communicationlink with electronics (e.g., the radar module 900) onboard the aerialplatform 100. As non-limiting examples, the communication link may beconfigured using Wi-Fi or WiMAX frequencies and protocols. A diversitycombiner 880 may be used to combine signals from multiple antennas.

A communication transceiver 870 (e.g., a Wi-Fi transceiver) may becoupled to the diversity combiner 880 and be configured to transmit andreceive frequencies to and from the diversity combiner 880. Acommunication modem 860 (e.g., a Wi-Fi modem) may be coupled to thecommunication transceiver 870 and be configured to package and modulatecommunication information for communication transmission as well asdemodulate and extract information from communication reception. Themicrocontroller 810 receives information from the communication modem860 and may perform operations related to the received information. Inaddition, based on processes performed on the microcontroller 810,information may be sent to the communication modem 860 for transmissionthrough the one or more antennas 890.

The microcontroller 810 may be coupled to a thrust controller 830, whichinterfaces with the alignment thrusters 622 and the divert thrusters 612(FIGS. 6A and 6B). A warhead fusing interface 840 may be provided tointerface to the warhead 440 (FIG. 4 ), the safe and arm module 460(FIG. 4 ) or a combination thereof, for arming and control of detonationof the warhead 440.

A roll sensor 850 and a vertical reference 855 may be used incombination to determine the attitude of the eject vehicle 400 as wellas a spin rate and spin position of the eject vehicle 400 andcommunicate such information to the microcontroller 810. Other types ofsensors, such as, for example, accelerometers and magnetometers may alsobe used for this purpose.

FIG. 9A is a block diagram illustrating elements that may be present onthe aerial platform 100 according to one or more embodiments of thepresent disclosure. The electronics module and functions thereof on theaerial platform 100 may be contained within a radar module 900, asillustrated in FIG. 9B. Alternatively, some of the function may bewithin the radar module 900 while other functions may be located indifferent places on the aerial platform 100 such as, for example, thecentral processor 360 (FIG. 3 ). The various modules used to control theradar module 900 and the eject vehicle 400 and determine otherinformation related thereto may be collectively referred to herein as an“onboard system.”

FIG. 9B is a perspective view of the radar module 900 that may bepresent on the aerial platform 100 according to one or more embodimentsof the present disclosure. The radar module 900 includes an azimuth scanradar antenna 920, an elevation scan radar antenna 940, and a wirelesscommunication link antenna 960.

The azimuth scan radar antenna 920 is included in an azimuth radarsubsystem, which includes a diplexer 922 for combining radar sent andreflected radar received. A radio frequency (RF) up/down converter 925converts the radar frequencies sent from a digital synthesizer 930 andconverts the radar frequencies received for use by a digital receiver935.

The elevation scan radar antenna 940 is included in an elevation radarsubsystem similar to the azimuth radar subsystem, but configured for theelevation direction. The elevation radar subsystem includes a diplexer942 for combining radar sent and reflected radar received. A radiofrequency (RF) up/down converter 945 converts the radar frequencies sentfrom a digital synthesizer 950 and converts the radar frequenciesreceived for use by a digital receiver 955.

The wireless communication link antenna 960 may be configured to providea communication link with electronics onboard the eject vehicle 400. Asnon-limiting examples, the communication link may be configured usingWi-Fi or WiMAX frequencies and protocols. A wireless communicationsubsystem includes a communication transceiver 965 (e.g., a Wi-Fitransceiver) coupled to the wireless communication link antenna 960 andconfigured to transmit and receive frequencies to and from the antenna960. A communication modem 970 (e.g., a Wi-Fi modem) may be coupled tothe communication transceiver 965 and be configured to package andmodulate communication information for communication transmission aswell as demodulate and extract information from communication reception.

A sector processor 910 communicates with the elevation radar subsystem,the azimuth radar subsystem, and the wireless communication subsystem.The sector processor 910 may communicate helicopter navigationinformation 912 from other electronics on the aerial platform 100.Referring also to FIG. 3 , the sector processor 910 may also communicatewith the dispenser 200 (e.g., one or more ALE-47s) using communicationsignal 914 and the missile approach warning system 320 (e.g., one ormore AAR-47s) using communication signal 916. The sector processor 910performs a number of functions to detect and track aerial threats 120,control and track the eject vehicle 400, as well as other functionsrelated to the active protection system. In some embodiments,communication between the dispenser 200 and the sector processor 910 maybe accomplished through the missile approach warning system 320.

The sector processor 910 in combination with the radar subsystems candetect and track incoming aerial threats 120 (e.g., RPGs). Based on thetracking of the incoming aerial threat 120, and in combination withnavigation information from the aerial platform 100, the sectorprocessor 910 can extrapolate to a geo-location of the attacker 110,from where the aerial threat 120 was launched. The aerial platform 100may act on this geo-location or transmit the geo-location to otheraerial platforms or ground-based platforms for follow-up actions.

The sector processor 910 may be configured to send launch commands tothe dispenser 200 on communication signal 914 to launch one or moreeject vehicles 400 to intercept one or more detected aerial threats 120.The sector processor 910 may also calculate required pitch adjustmentsthat should be performed by the eject vehicle 400 after it has beenejected and is safely away from the aerial platform 100.

Once the eject vehicle 400 is launched, the sector processor 910 may beconfigured to track the eject vehicle 400 and send guidance commands(i.e., divert commands) to the eject vehicle 400 so the eject vehicle400 can perform divert maneuvers to adjust its flight path toward theaerial threat 120. The sector processor 910 may also be configured todetermine when the eject vehicle 400 will be near enough to the aerialthreat 120 to destroy the aerial threat 120 by detonation of the warhead440 on the eject vehicle 400. Thus, a detonation command may be sent tothe eject vehicle 400 instructing it to detonate, or instructing it todetonate at a detonation time after receiving the command.

Referring to FIGS. 10A, 10B, 9, and 3 , the radar modules 900 may bemounted in close proximity to the existing AN/ALR-47 missile warningreceiver (MWR) installations to provide 360 degrees spatial coveragewhile minimizing wiring modifications to the helicopter. It isanticipated that an aerial threat 120 will be launched at relativelyshort ranges, typically on the order of 100 m. The radar modules 900 aredesigned to detect and track the low radar cross section (typically −15dBsm) of an RPG fired from any aspect angle, within 30 milliseconds oflaunch, and out to a range of at least 300 meters. The radars operate inthe Ka-Band to minimize the antenna size yet provide the precisionangular measurements needed to guide the eject vehicle 400 to interceptthe aerial threat 120. A high pulse-repetition-frequency pulse Dopplerwaveform provides radial velocity measurements as well as the clutterrejection needed to operate in close proximity to the ground whiledetecting low radar cross-section targets. Pulse compression may be usedto achieve precision range measurements as well as increasing thetransmit duty cycle to best utilize the capabilities of existing Ka-Bandsolid-state power amplifiers. The antennas generate a pair of orthogonalfan beams, providing a continuous track-while-scan capability tominimize detection latency and provide multiple target track capability.Beam scanning can be accomplished using a frequency scan method toeliminate the need for expensive phase shifters.

FIG. 10A illustrates an acquisition mode wherein an elevation radargenerates an elevation fan beam extending in the vertical direction thatsweeps in the horizontal direction and an azimuth radar generates anazimuth fan beam extending in the horizontal direction that sweeps inthe vertical direction. Thus, an entire 90-degree scan sector can becovered by the radar systems to quickly detect and acquire an incomingaerial threat 120 when it is within range.

FIG. 10B illustrates a track mode. In FIG. 10B, two sequential azimuthscans and two sequential elevation scans are shown that pinpoint a firstlocation 1010 of the eject vehicle 400. In addition, two sequentialazimuth scans and two sequential elevation scans are shown that pinpointa second location 1020 of the aerial threat 120. With this locationinformation, the sector processor 910 can derive relative positioninformation that can be used to provide divert commands to the ejectvehicle 400 to more closely intercept the aerial threat 120.

FIG. 11 is a spectrum diagram illustrating possible Doppler spectrumregions where various aerial vehicles may be detected. As non-limitingexamples, FIG. 11 illustrates a ground clutter spectrum 1110, a spectrum1120 for the eject vehicle 400 (i.e., PRJ in FIG. 11 ), a spectrum 1130that may be indicative of an RPG, and a spectrum 1140 that may beindicative of a MANPAD. Of course, other aerial threats and theirassociated spectrums may also be identified.

FIG. 12 is a simplified flow diagram illustrating some of the processes1200 involved in one or more embodiments of the present disclosure. Theprocesses may be loosely considered as an acquisition phase 1210, apre-launch phase 1220, an align and launch phase 1240, a guidance phase1260, a divert phase 1270, and a detonation phase 1280.

Operation block 1212 indicates that continuous radar scans are performedlooking for incoming aerial threats. Decision block 1214 indicates thatthe process loops until a target is detected. While not shown, duringthis phase the radar modules 900 may also be detecting distance andangle to wingman platforms (i.e., other aerial platforms) in thevicinity. Using communication between the various wingman platforms,sectors of responsibility can be identified as discussed more fullybelow in connection with FIG. 14 .

If a target is detected, the process 1200 enters the pre-launch phase1220. Operation block 1222 indicates that the sector processor 910 usesthe range and travel direction of the incoming aerial threat 120 tocalculate a threat direction to the incoming aerial threat 120 and anintercept vector pointing from a deployed eject vehicle 400 to aprojected intercept point where the eject vehicle 400 would interceptthe incoming aerial threat 120. Operation block 1224 indicates that theintercept vector is sent to the eject vehicle 400. The intercept vectormay be sent to the eject vehicle 400 in a number of forms. The actualdirectional coordinates may be sent and the eject vehicle 400 would beresponsible for determining the proper pitch maneuvers to perform.Alternatively, the sector processor 910 may determine the proper pitchmaneuvers that the eject vehicle 400 should perform after launch andsend only pitch commands (e.g., start and burn times for each alignmentthruster 622) to be used during the pitch maneuvers. While FIG. 12indicates that the intercept vector or pitch commands are sent beforelaunch, some embodiments may be configured such that this informationcan be sent after launch.

During the acquisition phase 1210 and pre-launch phase 1220, the ejectvehicle 400 remains in the dispenser 200 and connected to power. An RFcommunication link may be in operation through the eject vehicle 400antenna via a transmission line inside the dispenser 200.

The process enters the align and launch phase 1240 after the interceptvector is determined. Operation block 1242 indicates the impulsecartridge 750 is fired to propel the eject vehicle 400 from thedispenser 200 and safely away from the aerial platform 100.

Operation block 1244 indicates that the pitch maneuvers are performed toalign the eject vehicle 400 with the already determined interceptvector. The pitch maneuver is a two-stage process that sequentiallyexecutes an azimuth rotation and an elevation rotation to align thelongitudinal axis of the eject vehicle 400 along the intercept vector.The pitch maneuver does not have to be exact. As a non-limiting example,offsets of up to about 10 to 15 degrees may be corrected during flightof the eject vehicle 400 using the divert thrusters 612 during theguidance phase 1260. After ejection, the folding fins 482 will deployand the communication link antennas 960 will deploy and wirelesscommunication between the eject vehicle 400 and the radar module 900 maycommence.

Operation block 1246 indicates that the rocket motor 420 will fire,which accelerates the eject vehicle 400 to about 160 meters/second andimposes a spin rate on the eject vehicle 400 of about 10 Hertz. Uponexhaustion, the rocket motor 420 and folding fins 482 will separate andthe terminal vehicle (TV) is exposed. With separation of the TV, thesecond set of folding fins 484 deploy and the corner reflector 470 isexposed.

During the guidance phase 1260, the process will perform a track anddivert loop in order to adjust the flight path of the eject vehicle 400to more closely intercept the aerial threat 120. Operation block 1262indicates that the sector processor 910 will track the eject vehicle 400and aerial threat 120 as discussed above with reference to FIGS. 9A-10B.Decision block 1264, indicates that the sector processor 910 willdetermine if a divert maneuver is required to intercept the incomingaerial threat 120 and estimate the direction of divert thrust required.

A divert phase 1270 includes operations to cause the eject vehicle 400to modify its course. Operation block 1272 indicates that the divertdirection and time, if required, are sent to the eject vehicle 400.

The divert process takes into account the rotation of the eject vehicle400 and the direction of the desired divert thrust. This rotation adds acomplication to the selection and fire time determination of the properdivert thruster 612, but also ensures that all of the available divertthrusters 612 can be used to divert the eject vehicle 400 in any desireddirection substantially perpendicular to the travel direction of theeject vehicle 400. Operation block 1274 indicates that the processor onthe eject vehicle 400 will select the divert thruster 612 to be firedand determine the firing time based on the divert angle received fromthe sector processor 910 and its internal attitude sensors.

Operation block 1276 indicates that the appropriate divert thruster 612is fired at the appropriate fire time to move the eject vehicle 400laterally along a diversion vector to adjust the flight path of theeject vehicle 400. As a non-limiting example, each divert thruster 612may be capable of correcting for about two degrees of error from theinitial pointing of the eject vehicle 400 during the pitch maneuver.Thus, when the divert thrusters 612 are fired when the eject vehicle 400is in the correct rotational position, the process can slide the traveldirection vector of the eject vehicle 400 toward the path of the aerialthreat 120. Moreover, the process can fire in any circular direction andcan fire multiple divert thrusters 612 in the same direction torepeatedly move the eject vehicle 400 in the same direction.

While FIG. 12 indicates the guidance phase 1260 and the detonation phase1280 as operating sequentially, they also may operate in parallel.During the detonation phase 1280, operation block 1282 indicates thatthe sector processor 910 determines an optimum intercept time when theeject vehicle 400 will be at its closest point to the aerial threat 120.Operation block 1284 indicates that a detonation command may be sent tothe eject vehicle 400. This detonation command may be in the form of adetonation time for the eject vehicle 400 to count out or it may be inthe form of an immediate command for the eject vehicle 400 to perform assoon as the command is received.

Operation block 1286 indicates that the warhead 440 on the eject vehicle400 is detonated at the intercept time responsive to the detonationcommand received from the sector processor 910.

FIG. 13 illustrates an example flight path for the eject vehicle 400 andan aerial threat 120 during an intercept process. In this example, atypical RPG and EV trajectory example are shown. The RPG is launched ata range of about 100 meters and 30 degrees left of the nose of thehelicopter. The eject vehicle 400 receives its coordinate commands fromthe radar module 900 and is then ejected from the port chaff dispenser200 at an angle of 90 degrees to the helicopter axis.

During period 1310, the eject vehicle 400 separates to a distance ofabout two meters from the helicopter. During period 1320, the nosethrusters 622 pitch the eject vehicle 400 to the approximate approachangle of the incoming RPG (e.g., within about ±10° accuracy). The rocketmotor 420 then fires to accelerate the eject vehicle 400 toapproximately 160 meters/second and is then separated from the remainingterminal vehicle upon exhaustion.

During period 1330, the radar module 900 transmits a series of divertcommands to the eject vehicle 400, which fires the divert thrusters 612to correct the trajectory of the eject vehicle 400 and intercept theRPG. A radar command is finally sent to the eject vehicle 400 todetonate the warhead 440 when the terminal vehicle reaches the closestpoint of approach (CPA). The guidance algorithm may be configured toproduce a maximum CPA of about 30 centimeters, which is well within thelethal 0.6-meter kill radius of the warhead 440.

FIG. 14 illustrates two aerial vehicles flying in a formation andvarious radar sectors that may be covered by the aerial vehicles. Asignificant concern is the presence of wingman helicopters and thepotential damage caused by accidental targeting. The system presentedhas the capability of tracking and recognizing the adjacent helicoptersand networking with their associated active protection systems to avoidcollateral damage by handing off sectors covered by other platforms. InFIG. 14 , a first helicopter 1410 is monitoring a first radar sector1410A, a second radar sector 1410B, a third radar sector 1410C, and afourth radar sector 1410D.

A second helicopter 1420 near the first helicopter 1410 is monitoring afifth radar sector 1420A, a sixth radar sector 1420B, a seventh radarsector 1420C, and an eighth radar sector 1420D. If an aerial threatapproaches from a direction indicated by arrow 1430 it may be detectedby the third radar sector 1410C of the first helicopter 1410 and theseventh radar sector 1420C of the second helicopter 1420. If the firsthelicopter 1410 attempts to launch an eject vehicle, it may cause damageto the second helicopter 1420. However, using communication between thevarious wingman platforms, sectors of responsibility can be identified.Thus, for the direction indicated by arrow 1430, the first helicopter1410 can determine that the third radar sector 1410C will be covered bythe seventh radar sector 1420C of the second helicopter 1420. As aresult, while this formation continues, the first helicopter 1410 doesnot respond to threats in its third radar sector 1410C.

Returning to FIG. 9B, the radar module 900 may also be referred toherein more generically as an engagement management module (EMM) 900. Asdiscussed above, in some embodiments, an aerial platform 100 (FIGS. 1Aand 1B) may be configured with four engagement management modules 900,an example of which is shown in FIG. 3 .

The engagement management modules 900 may be used as part of ahelicopter active protection system (HAPS), but may also be used inother types of aerial vehicles, ground vehicles, water vehicles, andstationary deployments.

Returning to FIGS. 4 through 5C, the eject vehicle 400 may also bereferred to herein as an intercept vehicle 400 and a kill vehicle (KV)400. FIG. 5C illustrates the eject vehicle 400 after the rocket motor420 has burned and been detached from the airframe shell 430. At thisstage, the eject vehicle 400 may be referred to as a terminal vehicle, aterminal section, or a terminal section of the kill vehicle 400. Stillwithin the terminal section are the warhead 440, the divert thrustermodule 610, the alignment thruster module 620, the electronics module450 the battery 452, and the safe and arm module 460.

In operation, radar on the EMM 900 detects and tracks an aerial threat(e.g., an RPG) launched at the helicopter and launches one or more KVs400 from the AN/ALE-47 countermeasure dispenser to intercept theincoming RPG. Following launch, the KV 400 executes a series of pitchmaneuvers using nose thrusters (i.e., in the alignment thruster module620) to align the body axis with the estimated intercept point, uses aboost motor to accelerate to high speed to intercept the RPG at maximumrange, and executes a series of commands for lateral guidance maneuversusing a set of divert thrusters 610. Finally, the KV 400 warhead 440 iscommand detonated when the KV 400 reaches the closest point of approach(CPA) computed by an EMM guidance processor.

The flight time of the KV is typically on the order of 300 to 500 msec.During the short flight time, the KV is exposed to strong moments due todivert thruster offsets from the center of gravity and to strongaerodynamic moments due principally to the jet interaction (JI) effectwhen the divert thrusters are fired. The principle forces of interestare shown in FIG. 15 .

FIG. 15 is a simplified side view of a kill vehicle (KV) 400illustrating the principle forces of interest on the kill vehicle (KV)400. Ideally, the divert thruster's 610 center of thrust, X_(CT), wouldbe aligned with the KV 400 center of gravity (CG) to minimize anymoments introduced by its thrust force (F_(DT)). However, even withbalancing during manufacturing, there will be some migration of the CGas divert thrusters 610 and alignment thrusters 622 (FIGS. 6A and 6B andalso referred to herein as “pitch thrusters 622”) and “attitude controlthrusters 1604” (FIG. 16A) are fired and propellant is expended. Also,when the divert thrusters 610 are fired, airflow along a body 1510 ofthe KV 400 is disrupted, known as the jet interaction (JI) effect, whichintroduces a high pressure region 1506 in front of the divert thrusters610 and a low pressure region 1508 aft of the divert thrusters 610. Thiscauses a moment that can be represented as a force F_(JI) at locationX_(JI). The other principle aerodynamic component is the normal forceF_(N) introduced as a function of the angle of attack α, which is theangle between a body longitudinal axis 1520 and the velocity vector V.The normal force F_(N) acts at the center of pressure X_(CP).

One or more tail fins 1524 can be used to add aerodynamic stability, butthe time constants associated with these tail fins generally are notfast enough to stabilize the KV 400 during its short flight time.Instead, or in addition to the tail fins 1524, a plurality of smallmicro-thrusters 1604 (FIG. 16A) are added to the nose thruster module620 (FIGS. 4 and 16A) to implement an active attitude control. Thisintroduces F_(μ) at location X_(μ).

FIGS. 16A and 16B illustrate a nose thruster module 620. The nosethruster module 620 may include a plurality of pitch thrusters 622 and aplurality of attitude control thrusters 1604. FIG. 16A illustrates aplurality of cylinders 1606 and 1608 configured for storing propellantgrains. The larger cylinders 1606 correspond to the pitch thrusters 622while the smaller cylinders 1608 correspond to the attitude controlthrusters 1604. FIG. 16B illustrates a plurality of nozzles 1610 and1612, each configured to divert a flow from one of the plurality ofcylinders 1606, 1608. Thus, the combination of cylinders 1606 andnozzles 1610 create the pitch thrusters 622 and the combination ofcylinders 1608 and nozzles 1612 create the attitude control thrusters1604.

Referring again to FIG. 15 , the various moments described earlier maycause the KV 400 to become unstable and likely to tumble withoutstabilization. Maintaining the angle of attack α within a nominal 10 to20 degrees is useful to maintain the divert thrusters 612 normal to thevelocity vector V for proper guidance maneuvering and in directing thewarhead 440 blast toward the RPG at the point of detonation.

FIG. 17 is a simplified block diagram of an attitude control loop 1700.The EMM 900 radar generates position {circumflex over (d)}_(KV) andvelocity {circumflex over (ν)}_(KV) track data for the KV 400 and sendsthis information to the KV 400 via a command link radio (CLR) 1706. TheEMM 900 radar also transmits a roll angle {circumflex over (ϕ)}_(TX) viaa transmit antenna 1804 (FIG. 18 ) to the CLR 1706 to the KV 400 overthe CLR 1706. A joint adaptive polarization and roll angle estimator(JAPRAE) algorithm 1710, resident in the KV 400, processes the polarizedCLR 1706 signal received by two orthogonal linear receive antennas 1802(FIG. 18 ) on the KV 400 to estimate a roll angle {circumflex over(ϕ)}_(KV) of the KV 400. An onboard inertial management unit (IMU) 1714measures an acceleration a_(KV) and an angular velocity ω_(KV) of the KV400. This information is processed by an extended Kalman filter (EKF)1720 to generate fused estimates of the position {tilde over (d)}_(KV)velocity {tilde over (ν)}_(KV), attitude [{tilde over (ϕ)}_(KV), {tildeover (θ)}_(KV), {tilde over (ψ)}_(KV)], and angular velocity {tilde over(ω)}_(KV) of the KV 400.

The attitude control algorithm 1730 is designed to fire the nose mountedattitude control thrusters 1604 to maintain the angle of attack α (FIG.15 ) within preset limits. Once it is determined that an attitudecorrection thrust is required, the attitude control algorithm 1730selects the next available attitude control thruster 1604 that comesinto position due to the body spin at operational block 1732, and timesthe firing to coincide with the desired direction of thrust atoperational block 1734. The attitude control algorithm 1730 maintainsknowledge of which attitude control thrusters 1604 have been fired andwhich are available at operational block 1736.

The KV 400 has an onboard IMU 1714 that measures the angular rate andacceleration. The IMU 1714 may include a set of 3-axis gyros, 3-axisaccelerometers, 3-axis magnetometers, and a field programmable gatearray (FPGA) to partially process the sensor information. The sensorsmay be solid-state MEMS that are machined on a small circuit board. TheIMU 1714 also includes an FPGA that contains the EKF 1720 to fuse theoutput of the sensors into 3-dimensional position, velocity, angularvelocity and attitude.

FIG. 18 is a simplified flowchart of the joint adaptive polarization androll angle estimator (JAPRAE) algorithm 1710. The JAPRAE algorithm 1710may process the CLR 1706 signals received, for example, by orthogonallinearly polarized receive antennas 1802, to adapt the two receivesignals in a beam former 1806 (also referred to herein as “the dualpolarized receiver 1806”) to match the incoming polarization toeliminate polarization mismatch loss induced by body rotation. TheJAPRAE algorithm 1710 also processes the CLR 1706 signals received toprovide an estimate of the KV 400 roll angle {circumflex over (ϕ)}_(rx).

The electric vector field of a propagating signal can be represented bya pair of orthogonal components:

$\begin{matrix}{{E_{tx}(t)} = {{{\lbrack {u_{tx\phi 1},u_{tx\phi 2}} \rbrack\begin{bmatrix}{p_{1}(t)} \\{p_{2}(t)}\end{bmatrix}}E_{tx}e^{i\omega_{0}t}} = {U_{tx}pE_{tx}e^{i\omega_{0}t}}}} & (1)\end{matrix}$

where u_(txϕ1) and u_(txϕ2) are orthogonal unit vectors that define thedirection of the electric field components, p_(1hu (t)) and p₂ ^((t))are the complex components of a unit vector that define the projectionof the electric field vector on the basis vectors, and U and p are thematrix of basis unit vectors and the polarization vector, respectively.Similarly, the polarization of an antenna can be defined by a 2×1complex vector q. The signal received by an antenna with polarization qfrom a signal with polarization p is simply the inner product of p andq:

x _(rx)(t)=q(t)^(H) p(t)E(t)e ^(iβ(t))  (2)

If polarizations p and q are defined in different coordinate systems,U_(tx) and U_(rx), then (2) becomes:

x _(rx)(t)=q(t)^(H) U _(rx) ^(T) U _(tx) p(t)E(t)e ^(iβ(t))  (3)

Assume the polarization of a source antenna 1804 and the receiveantennas 1802 can be represented as p_(o) and q_(o) in terms of anatural antenna coordinate system. Let A_(tx)(t) and A_(rx) (t) be thetime varying coordinate transformations from the natural coordinatesystem to some inertial system. In this system the electric field vectorand receiver polarization vectors are given by:

E _(tx)(t)=A _(tx)(t)U _(tx) p _(o) E(t)e ^(iβ(t))

q(t)=A _(rx)(t)U _(rx) q _(o).  (4)

Then, (3) becomes:

x _(rx)(t)=q _(o) ^(H) U _(rx) ^(T) A _(rx) ^(T)(t)A _(tx)(t)U _(tx) p_(p) E(t)e ^(iβ(t))  (5)

In the JAPRAE system 1710 there are two orthogonal receive antennas1802, with polarizations q_(o1) and q_(o2). Thus, the signal generatedat the input of the dual polarized receiver 1806 is given by:

$\begin{matrix}\begin{matrix}{{x_{rx}(t)} = {\begin{bmatrix}{x_{rx1}(t)} \\{x_{rx2}(t)}\end{bmatrix} = {\begin{bmatrix}{q_{o1}^{H}U_{rx}^{T}{A_{rx}^{T}( \phi_{rx} )}{A_{tx}( \phi_{tx} )}U_{tx}p_{o}} \\{q_{o2}^{H}U_{rx}^{T}{A_{rx}^{T}( \phi_{rx} )}{A_{tx}( \phi_{tx} )}U_{tx}p_{o}}\end{bmatrix}{E(t)}e^{i{\beta(t)}}}}} \\{= {{\begin{bmatrix}a_{1} \\a_{2}\end{bmatrix}{E(t)}e^{i{\beta(t)}}} = {a{E(t)}e^{i{\beta(t)}}}}}\end{matrix} & (6)\end{matrix}$

In equation (6) the transformations are explicitly represented asrotations of the transmitter ϕ_(tx) and receiver ϕ_(rx) around the lineof sight. In the case where the transmit antenna 1804 and the receiveantennas 1802 are linearly polarized, it can be shown that this reducesto the following form:

$\begin{matrix}{{x_{rx}(t)} = {{\begin{bmatrix}{\cos( {\phi_{tx} - \phi_{rx}} )} \\{- {\sin( {\phi_{tx} - \phi_{rx}} )}}\end{bmatrix}{E(t)}e^{i{\beta(t)}}} = {\begin{bmatrix}{\cos(\delta)} \\{- {\sin(\delta)}}\end{bmatrix}{E(t)}e^{i{\beta(t)}}}}} & (7)\end{matrix}$

Consider a signal y(t) generated by performing a rotation on the vectorx(t).

$\begin{matrix}\begin{matrix}{{y_{rx}(t)} = {\begin{bmatrix}{y_{rx1}(t)} \\{y_{rx2}(t)}\end{bmatrix} = {{\begin{bmatrix}{\cos( \hat{\delta} )} & {- {\sin( \hat{\delta} )}} \\{\sin( \hat{\delta} )} & {\cos( \hat{\delta} )}\end{bmatrix}\begin{bmatrix}{\cos(\delta)} \\{- {\sin(\delta)}}\end{bmatrix}}{E(t)}e^{i{\beta(t)}}}}} \\{= {\begin{bmatrix}{{{\cos( \hat{\delta} )}{\cos(\delta)}} + {{\sin( \hat{\delta} )}{\sin(\delta)}}} \\{{{\sin( \hat{\delta} )}{\cos(\delta)}} - {{\cos( \hat{\delta} )}{\sin(\delta)}}}\end{bmatrix}{E(t)}e^{i{\beta(t)}}}} \\{= {\begin{bmatrix}{\cos( {\hat{\delta} - \delta} )} \\{\sin( {\hat{\delta} - \delta} )}\end{bmatrix}{E(t)}e^{i{\beta(t)}}}}\end{matrix} & (8)\end{matrix}$

Note that if {circumflex over (δ)}=δ, y_(rx1)=Ee^(iβ(t)) and y_(rx0)=0.This amounts to a case where γ_(rx1) is the output of a beamformer thatmatches the polarization of the incident signal to maximize the receivesignal power. It can be shown that this also maximizes thesignal-to-noise ratio.

Now form the ratio of y2(k) to y1(k)

$\begin{matrix}\begin{matrix}{{r(t)} = {\frac{y_{{rx}2}(t)}{y_{{rx}1}(t)} = {\frac{{\sin( {\hat{\delta} - \delta} )}{E(t)}e^{i{\beta(t)}}}{{\cos( {\hat{\delta} - \delta} )}{E(t)}e^{i{\beta(t)}}} = {\tan( {\hat{\delta} - \delta} )}}}} \\{= {\tan( {\hat{\delta} - \delta} )}}\end{matrix} & (9)\end{matrix}$

This forms the basis for generating an error signal for an adaptivepolarization loop.

$\begin{matrix}{{e_{\delta}(t)} = {{{\hat{\delta}(t)} - {\delta(t)}} = {\tan^{- 1}( \frac{y_{2}(t)}{y_{1}(t)} )}}} & (10)\end{matrix}$

FIG. 19 illustrates the JAPRAE algorithm 1710 (FIG. 17 ) in the form ofa conventional servo loop 1900. The closed loop may force y_(ex2) tozero, hence {circumflex over (δ)}→δ. Under closed loop conditions, theKV 400 roll angle is estimated by:

{circumflex over (θ)}_(rx)(t)={circumflex over (θ)}_(tx)(t)−{circumflexover (δ)}(t)  (11)

The transmitter roll angle, {circumflex over (ϕ)}_(tx) can betransmitted to the KV 400 via the data link 1806.

FIG. 20 illustrates a non-limiting example of a set of simulationresults 2000 of the JAPRAE algorithm 1710 (FIG. 17 ) tracking the rollangle of a KV 400 body spinning on its longitudinal axis 1520 (FIG. 15 )at 2200 degrees per second. A top plot 2002 of FIG. 20 illustrates anoverlay of true and estimated roll angles. A bottom plot 2004illustrates a difference or error between the overlay of true andestimated roll angles. The set of simulation results 2000 are shownwithout additive noise and illustrate the performance of the servo loop1900 (FIG. 19 ).

Note that the solution for equation 10 is ambiguous by n radians.Therefore, the loop 1900 of FIG. 19 should be initialized near thecorrect angle to avoid the ambiguity.

A Kalman filter methodology is used to fuse the sensor data. Thetranslational motion involves a set of linear equations and theconventional Kalman filter (KF) can be used, whereas the rotationalmotion involves a set of nonlinear equations and the conventionalextended Kalman filter (EKF) can be used.

A 9×1 discrete time translational motion state vector, x(k), includes3-D spatial components comprising a position vector, d(k), a velocityvector, v(k), and an acceleration vector, a(k), as follows:

x(k)=[d ^(T)(k),v ^(T)(k),a ^(T)(k)]^(T)  (12)

Transition and measurement equations include a transition matrix, F_(k),a plant noise coupling matrix, G_(k), a measurement matrix, H_(k), statetransition noise w_(k) and measurement noise V_(k).

x _(k+1) =Fx _(k) +G _(k) w _(k)

Z _(k) =H _(k) x _(k) +V _(k)  (13)

The state transition noise w_(k) is i.i.d. zero mean Gaussian withcovariance R_(k), and the measurement noise V_(k) is also i.i.d. zeromean Gaussian with covariance Q_(k). The subscript k indicates thevariables might be time varying.

The expanded form of these equations is given by:

$\begin{matrix}{{\begin{bmatrix}d_{k + 1} \\v_{k + 1} \\a_{k + 1}\end{bmatrix} = {{\begin{bmatrix}I_{3} & {T_{s}I_{3}} & {{0.5}T_{s}^{2}I_{3}} \\0_{3} & I_{3} & {T_{s}I_{3}} \\0_{3} & 0_{3} & I_{3}\end{bmatrix}\begin{bmatrix}d_{k} \\v_{k} \\a_{k}\end{bmatrix}} + {\begin{bmatrix}{{0.5}T_{s}^{2}I_{3}} \\{T_{s}I_{3}} \\I_{3}\end{bmatrix}\begin{bmatrix}w_{d,k} \\w_{v,k} \\w_{{wa},k}\end{bmatrix}}}}{ \lbrack\begin{matrix}z_{d,k} \\z_{v,k} \\z_{a,k}\end{matrix} \rbrack = {{\begin{bmatrix}I_{3} & 0_{3} & 0_{3} \\0_{3} & I_{3} & 0_{3} \\0_{3} & 0_{3} & I_{3}\end{bmatrix}\begin{bmatrix}d_{k} \\v_{k} \\a_{k}\end{bmatrix}} + \begin{bmatrix}v_{d,k} \\v_{v,k} \\v_{a,k}\end{bmatrix}}}} & (14)\end{matrix}$

where T_(s) is the sample interval and I₃ and 0 ₃ are 3×3 unity and zeromatrices, respectively.

The Kalman filter 1720 (FIG. 17 ) involves two acts: the projection actand the update act.

The projection act is given by:

{circumflex over (x)} _(k|k+1) =F _(k) {circumflex over (x)} _(k−1|k−1)

{circumflex over (P)} _(k|1+1) =F _(k) {circumflex over (P)} _(k−1|k−1)F _(k) ^(T) +G _(k) Q _(k) G _(k) ^(T)  (15)

The update act is given by:

L _(k) =P _(k|k−1) H _(k) ^(T)[H _(k) {circumflex over (P)} _(k|k−1) H_(k) ^(T) +R _(k)]⁻¹

{circumflex over (x)} _(k|k) =F{circumflex over (x)} _(k|k−1) +L _(k)[Z_(k) −H _(k) {circumflex over (x)} _(k|k−1)]

{circumflex over (P)} _(k|k) ={circumflex over (P)} _(k|k−1)−{circumflex over (P)} _(k|k−1) H _(k) ^(T)[H _(k) {circumflex over (P)}_(k|k−1) H _(k) ^(T) +R _(k)]⁻¹ H _(k) ^(T) {circumflex over (P)}_(k|k−1) ^(T)  (16)

The 9×1 discrete time translational motion state vector, x(k) comprises3-D spatial components including attitude, θ(k), and angle rate, ω(k).An overbar is used to distinguish the rotational variables from thetranslational variables:

x (k)=[θ^(T)(k),ω^(T)(k)]^(T)  (17)

The transition and measurement equations have a similar form as thetranslational equations except the transition equation is nonlinear.

x _(k+1) =F _(k)(y _(k))+ G _(k) w _(k)

Z _(k+1) =H _(k) y _(k)+ν _(k)  (18)

The transition noise, w _(k) is i.i.d. zero mean Gaussian withcovariance R _(k) and the measurement noise v _(k) is i.i.d. Gaussianwith covariance Q _(k). The nonlinear property is shown in the expandedform.

$\begin{matrix}{{\begin{bmatrix}{\overset{¯}{x}}_{\theta,{k + 1}} \\{\overset{¯}{x}}_{\omega,{k + 1}}\end{bmatrix} = {{\begin{bmatrix}I_{3} & {T_{s}{B( \theta_{k} )}} \\0_{3} & I_{3}\end{bmatrix}\begin{bmatrix}{\overset{\_}{x}}_{\theta,k} \\{\overset{\_}{x}}_{\omega,k}\end{bmatrix}} + {\begin{bmatrix}0_{3} \\I_{3}\end{bmatrix}\begin{bmatrix}{\overset{\_}{w}}_{\theta,k} \\{\overset{\_}{w}}_{\omega,k}\end{bmatrix}}}}{ \lbrack\begin{matrix}{\overset{\_}{z}}_{\phi,k} \\{\overset{\_}{z}}_{\omega,k}\end{matrix} \rbrack = {{\begin{bmatrix}1 & 0 & 0 & 0 & 0 & 0 \\ & 0_{3} & & & I_{3} & \end{bmatrix}\begin{bmatrix}{\overset{¯}{x}}_{\theta,k} \\{\overset{¯}{x}}_{\omega,k}\end{bmatrix}} + \begin{bmatrix}{\overset{\_}{v}}_{\phi,k} \\{\overset{\_}{v}}_{\omega,k}\end{bmatrix}}}} & (19)\end{matrix}$

The nonlinear matrix B(θ_(k)) converts angle rate to Euler angle rateand is given by:

$\begin{matrix}{{B( \theta_{k} )} = \begin{bmatrix}1 & {{\sin(\phi)}{\tan(\theta)}} & {{\cos(\phi)}{\tan(\theta)}} \\0 & {\cos(\phi)} & {- {\sin(\phi)}} \\0 & {{\sin(\phi)}{\sec(\theta)}} & {{\cos(\phi)}{\sec(\theta)}}\end{bmatrix}} & (20)\end{matrix}$

The EKF method is similar to the KF method except the transitionequation is linearized using the Jacobian.

$\begin{matrix}{{\overset{\_}{F}}_{k} = {\frac{\partial{{\overset{\_}{F}}_{k}( y_{k} )}}{\partial{\overset{\_}{x}}_{k}} = \lbrack {\frac{\partial{{\overset{\_}{F}}_{k}( y_{k} )}}{\partial\theta}\frac{\partial{{\overset{\_}{F}}_{k}( y_{k} )}}{\partial\omega}} \rbrack}} & (21)\end{matrix}$

The two derivatives are:

${{\frac{\partial{{\overset{\_}{F}}_{k}( y_{k} )}}{\partial\theta} = \begin{bmatrix}{1 + {T_{s}( {{{\cos(\phi)}{\tan(\theta)}\omega_{y,k}} - {{\sin(\phi)}{\tan(\theta)}\omega_{z,k}}} )}} & {T_{s}( {{{\sin(\phi)}{\sec^{2}(\theta)}\omega_{y,k}} + {{\cos(\phi)}{\sec^{2}(\theta)}\omega_{z,k}}} )} & 0 \\{{{- T_{s}}{\sin(\phi)}\omega_{y,k}} - {T_{s}{\cos(\phi)}\omega_{z,k}}} & 1 & 0 \\{T_{s}( {{{\cos(\phi)}{\sec(\theta)}\omega_{y,k}} - {{\sin(\phi)}{\sec(\theta)}\omega_{z,k}}} )} & {T_{s}( {{{\sin(\phi)}{\sec(\theta)}{\tan(\theta)}\omega_{y,k}} + {{\cos(\phi)}{\sec(\theta)}{\tan(\theta)}\omega_{z,k}}} )} & 1\end{bmatrix}}\text{⁠}}{\frac{\partial{{\overset{\_}{F}}_{k}( y_{k} )}}{\partial\omega} = {T_{s}{B( \theta_{k} )}}}$

From this point on, the EKF method is similar to the KV method using thelinearized components.

The projection step for the EKF is given by:

{circumflex over (x)} _(k|k+1) =F _(k) {circumflex over (x)} _(k−1|k−1)

{circumflex over (P)} _(k|k+1) =F _(k) {circumflex over (P)} _(k−1|k−1)+G _(k) Q _(k) G _(k) ^(T)  (22)

The update step for the EKF is given by:

L _(k)= {circumflex over (P)} _(k|k−1) H _(k) ^(T)[H _(k) {circumflexover (P)} _(k|k−1) H _(k) ^(T) +R _(k)]⁻¹

{circumflex over (x)} _(k|k) =F {circumflex over (x)} _(k|k−1) +L _(k)[Z_(k) −H _(k) {circumflex over (x)} _(k|k−1)]

{circumflex over (P)} _(k|k)= {circumflex over (P)} _(k|k−1)−{circumflex over (P)} _(k|k−1) H _(k) ^(T)[H _(k) {circumflex over (P)}_(k|k−1) H _(k) ^(T) +R _(k)]⁻¹ H _(k) ^(T) {circumflex over (P)}_(k|k−1) ^(T)  (23)

FIGS. 21, 22, and 23 illustrate sets of simulation results of the EKFfusion [0163] algorithm using the EMM 900 radar, JAPRAE 1710 and IMU1714 (FIG. 17 ) inputs for a typical KV 400 flight profile. FIG. 21illustrates position results of the simulation. A top plot illustratestrue position values, a middle plot illustrates estimated positionvalues, and a bottom plot illustrates position error. An x position isindicated by solid lines or blue lines, a y position is indicated bydotted lines or green lines, and a z position is indicated by dashedlines or red lines.

FIG. 22 illustrates velocity results of the simulation. A top plotillustrates true velocity values, a middle plot illustrates estimatedvelocity values, and a bottom plot illustrates velocity error. An xvelocity is indicated by solid lines or blue lines, a y velocity isindicated by dotted lines or green lines, and a z velocity is indicatedby dashed lines or red lines.

FIG. 23 illustrates attitude results of the simulation. A top plotillustrates true attitude values, a middle plot illustrates estimatedattitude values, and a bottom plot illustrates attitude error. An xattitude is indicated by solid lines or blue lines, a y attitude isindicated by dotted lines or green lines, and a z attitude is indicatedby dashed lines or red lines.

FIG. 24 illustrates an attitude control loop 2400. The attitude controlloop 2400 utilizes an uplink and onboard EKF outputs to compute firingcommands to the plurality of attitude control thrusters (ACT) 1604 tomaintain the KV 400 body axis 1520 aligned with the velocity vector V(FIG. 15 ).

In FIG. 24 , {tilde over (d)}_(KV) and {tilde over (ν)}_(KV) areprojectile position and velocity vectors in the north-east-down (NED)frame. [{tilde over (ϕ)}_(KV), {tilde over (θ)}_(KV), {tilde over(ψ)}_(KV)] are body Euler angles and {tilde over (ω)}_(KV) is body ratevector. The projectile mass, CG location and inertia tensor are updatedwhenever there is a divert thruster 610 or an attitude control thruster1604 firing. The attitude control loop 2400 may include an attitudecontrol algorithm 2402. The attitude control algorithm 2402 may computehow much attitude control thrust is needed at the attitude thrusterstation 620 and which direction the thrust centroid should point to. Ifthe attitude control thrust (ACT) exceeds a certain fraction of the ACTforce, then the algorithm may initiate a command to fire an ACT. Theattitude thruster select function 2404 then searches for an ACT to fireat operational block 2404. If the selected ACT is at the desiredignition position, the attitude thruster fire control 2406 electronicssquib the selected ACT to fire at operational block 2406. Otherwise, theprocess exits the control loop. Finally, the attitude thruster unitsremaining function 2410 keeps track of which ACTs are spent and whichare available for use. This information is fed back to the mass propertyestimator function 2412 and the attitude thruster select function 2404.

Operation may be as explained in acts 1 through 12 below:

Act 1: The onboard EKF algorithm tracks the projectile yaw ({tilde over(ψ)}_(KV)), pitch ({tilde over (θ)}_(KV)) and roll ({tilde over(ϕ)}_(KV)) attitude angles. These angles are used to compute the NED tobody directional cosine matrix C_(n) ^(b) as

$\begin{matrix}{C_{n}^{b} = {\lbrack {\begin{matrix}1 & 0 & 0 \\0 & {\cos{\overset{\sim}{\phi}}_{KV}} & {\sin{\overset{\sim}{\phi}}_{KV}} \\0 & {{- \sin}{\overset{\sim}{\phi}}_{KV}} & {\cos{\overset{\sim}{\phi}}_{KV}}\end{matrix}} \rbrack \cdot \lbrack {\begin{matrix}{\cos{\overset{\sim}{\theta}}_{KV}} & 0 & {{- \sin}{\overset{\sim}{\theta}}_{KV}} \\0 & 1 & 0 \\{\sin{\overset{\sim}{\theta}}_{KV}} & 0 & {\cos{\overset{\sim}{\theta}}_{KV}}\end{matrix}} \rbrack \cdot \lbrack {\text{⁠}\begin{matrix}{\cos{\overset{\sim}{\psi}}_{KV}} & {\sin{\overset{\sim}{\psi}}_{KV}} & 0 \\{{- \sin}{\overset{\sim}{\psi}}_{KV}} & {\cos{\overset{\sim}{\psi}}_{KV}} & 0 \\0 & 0 & 1\end{matrix}} \rbrack}} & (24)\end{matrix}$

Act 2: The projectile velocity in the NED frame, {tilde over (V)}_(KV),is tracked by EMM 900 radar and uplinked to the projectile via RFcommunication links. The velocity is then transformed into the bodyframe.

{tilde over (V)} _(KV) ^(b) =C _(n) ^(b) ·{tilde over (V)} _(KV)  (25)

Act 3: The total angle of attack α_(total) and aerodynamic roll anglesϕ_(aero) is then computed. The total angle of attack should be small atall times. The aerodynamic angle tells where the total angle of attackis relative to the body roll axis.

$\begin{matrix}{{\alpha_{total} = {\tan^{- 1}( \frac{\sqrt{{{\overset{\sim}{V}}_{KV}^{b}(2)}^{2} + {{\overset{\sim}{V}}_{KV}^{b}(3)}^{2}}}{{\overset{\sim}{V}}_{KV}^{b}(1)} )}}{\varphi_{aero} = {\tan^{- 1}( \frac{{\overset{\sim}{V}}_{KV}^{b}(2)}{{\overset{\sim}{V}}_{KV}^{b}(3)} )}}} & (26)\end{matrix}$

Act 4: The attitude changing rate cannot be measured for the purpose ofrate feedback, but the synthetic rates may be derived by using therelationship between the attitude rate and the estimated body rates andattitude angles as

{dot over (ω)}_(KV)=({tilde over (ω)}_(KV)(2)sin {tilde over(ϕ)}_(KV)+{tilde over (ω)}_(KV)(3)cos {tilde over (ϕ)}_(KV))sec {tildeover (θ)}_(KV)

{dot over (θ)}_(KV)={tilde over (ω)}_(KV)(2)cos {tilde over(ϕ)}_(KV)−{tilde over (ω)}_(KV)(3)sin {tilde over (ϕ)}_(KV)

{dot over (ϕ)}_(KV)={tilde over (ω)}_(KV)(1)+{dot over (ψ)}_(KV) sin{tilde over (ϕ)}_(KV)  (27)

Act 5: Compute the KV 400 heading and flight path angles in the NEDframe. These are the desired body yaw and pitch attitude angles that wewant to control. Note that the body roll angle is free.

$\begin{matrix}{{\psi_{cmd} = {\tan^{- 1}( \frac{{\overset{\sim}{V}}_{KV}(2)}{{\overset{\sim}{V}}_{KV}(1)} )}}{\theta_{cmd} = {\tan^{- 1}( \frac{- {{\overset{\sim}{V}}_{KV}(3)}}{\sqrt{{{\overset{\sim}{V}}_{KV}(1)}^{2} + {{\overset{\sim}{V}}_{KV}(2)}^{2}}} )}}} & (28)\end{matrix}$

Act 6: Compute attitude error by subtracting the estimated attitudeangles from the commanded attitude angles above. This error is thenmultiplied by a proportional gain K_(p,outer) to form attitude anglerate command.

{dot over (ψ)}_(cmd) =K _(p,outer)(ψ_(cmd)−{tilde over (ψ)}_(KV))

{dot over (θ)}_(cmd) =K _(p,outer)(θ_(cmd)−{tilde over (θ)}_(KV))  (29)

Act 7: Compute attitude rate errors by subtracting the attitude ratederived in Act 4 above from the rate commands above. These rate errorsare multiplied by a proportional gain K_(p,inner) to form attitudeacceleration commands.

{umlaut over (ψ)}_(cmd) =K _(p,inner)({dot over (ψ)}_(cmd)−{dot over(ψ)}_(KV))

{umlaut over (θ)}_(cmd) =K _(p,inner)({dot over (θ)}_(cmd)−{dot over(θ)}_(KV))  (30)

FIG. 25 illustrates a controller input-output topology. The yaw andpitch commands are derived from Act 5. The estimated body attitudeangles, Euler, and body rates, ω, are provided by the JAPRAE algorithm1710. The controller gain Kp_(outer) and Kp_(inner) are to be andoptimized for various engagement scenarios. The symbol Euler_(d)represents the Euler rate vector └{dot over (ψ)}_(KV), {dot over(θ)}_(KV), {dot over (θ)}_(KV)┘. Euler_(dd) represents the Euleracceleration vector {umlaut over (└)}_(Kv), {umlaut over (θ)}_(KV),{umlaut over (ϕ)}_(KV)┘. F_(cmd) is the output of the controllercontaining the magnitude of the attitude thrust force command and itsorientation relative to the body roll axis 1520. The function thatcomputes the force command is described in Acts 8 through 10.

Act 8: Because the air vehicle is aerodynamically unstable at all flightregimes, angle of attack will grow if unchecked. For this reason, theaerodynamic moment to be countered must be estimated, in addition to themoment needed to produce the attitude acceleration of Act 7. Theestimation is done by storing a complete set of aerodynamic coefficienttables on board the projectile and computing the total yawing andpitching moments about the estimated CG location on the flight. Thealgorithm to estimate the KV 400 mass, CG location and inertia tensorwill be described below with respect to FIG. 27 . The total aerodynamicmoment in the body frame, ignoring the rolling moment, may be computedas:

$\begin{matrix}{{M^{b} = {{QSd}\begin{Bmatrix}0 \\C_{m} \\C_{n}\end{Bmatrix}}}{C_{m} = {{( {{\overset{\_}{X}}_{CP} - {\overset{\_}{X}}_{CG}} )*C_{N}} + {C_{m_{q}}\overset{\_}{q}} + {C_{A}{\overset{\_}{Z}}_{CG}}}}{C_{n} = {{( {{\overset{\_}{X}}_{CP} - {\overset{\_}{X}}_{CG}} )*C_{Y}} + {C_{n_{r}}\overset{\_}{r}} - {C_{A}{\overset{\_}{Y}}_{CG}}}}} & (31)\end{matrix}$

Where:

Q is dynamic pressure

S is reference area

d is reference diameter

C_(m) is total pitching moment coefficient

C_(n) is total yawing moment coefficient)

X _(CP) is normalized center of pressure location along the body x-axis)

X _(CG) is normalized center of gravity location along the body x-axis

Y _(CG) is CG offset from the body x-axis

Z _(CG) is CG offset from the body x-axis

C_(A) is axial drag-force coefficient

C_(N) is normal force coefficient

C_(Y) is lateral force coefficient

C_(m) _(g) is pitch damping coefficient

C_(n), is yaw damping coefficient

q is normalized pitching rate

r is normalized yawing rate.

Act 9: The instantaneous moment about the CG location required toproduce the attitude acceleration which will force the body to alignwith the velocity vector V (FIG. 15 ) can now be calculated as:

$\begin{matrix}{M_{cmd} = {{- \begin{bmatrix}0 & 0 & 0 \\0 & {{- \sin}{\overset{\sim}{\phi}}_{KV}} & {{- \cos}{\overset{\sim}{\phi}}_{KV}} \\0 & {{- \cos}{\overset{\sim}{\phi}}_{KV}} & {\sin{\overset{\sim}{\phi}}_{KV}}\end{bmatrix}}\begin{Bmatrix}0 \\c_{2} \\c_{3}\end{Bmatrix}{I_{t}.}}} & (32)\end{matrix}$

I_(t) is the estimated moment of inertia about the lateral axis and, C₂and C₃ are computed as:

c ₂=COS {tilde over (θ)}_(KV){umlaut over (ψ)}_(cmd)−sin {tilde over(θ)}_(KV){dot over (θ)}_(KV){dot over (ψ)}_(KV)−({tilde over(ω)}_(KV)(2)cos {tilde over (ϕ)}_(KV)−{tilde over (ω)}_(KV)(3)sin {tildeover (ϕ)}_(KV)){dot over (ψ)}_(KV)−

(I _(t) −I _(a))({tilde over (ω)}_(KV)(1){tilde over (ω)}_(KV)(3)sin{tilde over (ϕ)}_(KV)+{tilde over (ω)}_(KV)(1){tilde over(ω)}_(KV)(2)cos {tilde over (ϕ)}_(KV))

C ₃={umlaut over (θ)}_(cmd)+({tilde over (ω)}_(KV)(2)sin {tilde over(ϕ)}_(KV)+{tilde over (ω)}_(KV)(3)cos {tilde over (ϕ)}_(KV)){dot over(ϕ)}_(KV)−

(I _(t) −I _(a))({tilde over (ω)}_(KV)(1){tilde over (ω)}_(KV)(3)cos{tilde over (ϕ)}_(KV)−{tilde over (ω)}_(KV)(1){tilde over (ω)}KV(2)Sin{tilde over (ϕ)}_(KV)  (33)

I_(a) denotes the estimated moment of inertia about the roll axis.

Act 10: Given the estimated CG location, X_(CG) and the attitude nozzlelocation, X_(AT), we can convert the above moment command to thrustforce command as:

$\begin{matrix}{F_{cmd} = {\begin{Bmatrix}0 \\{{M_{cmd}(3)} - {M^{b}(3)}} \\{{- {M_{cmd}(2)}} + {M^{b}(2)}}\end{Bmatrix}\frac{1}{X_{AT} - X_{CG}}}} & (34)\end{matrix}$

The magnitude of this command is:

F _(total) =∥F _(cmd)∥  (35)

And, the roll orientation relative to the body y-axis is:

$\begin{matrix}{\phi_{cmd} = {\tan^{- 1}( \frac{F_{cmd}(3)}{F_{cmd}(2)} )}} & (36)\end{matrix}$

Act 11: If the thrust force command, F_(total), is greater than athreshold value, usually a fraction of the attitude thruster force, athruster search algorithm may be initiated to find a thruster to fire.Otherwise, no attitude thruster may be fired.

FIG. 26 illustrates an example of a search procedure that may be used tofind a thruster to fire. A plurality of attitude thrusters 1, 2, 3, 4,5, 6, 7, and 8 are shown with circles. A white circle, such as attitudethrusters 1, 4, 6, and 8, indicates the attitude thruster has beenspent. A shaded circle, such as attitude thrusters 2, 3, 5, and 7,indicates the attitude thruster has not been spent.

From Act 3, a total angle of attack is found between the body x-axis andthe velocity vector V^(b) and the aerodynamic angle relative to the bodyz-axis. To reduce the total angle of attack, an attitude thruster may befired 180 degrees opposite to ϕ_(aero) such that the resultant force Fpushes the body nose toward the velocity vector. Knowing the duration ofthe attitude thruster burn and time delays from the thruster commandgeneration to actual thruster ignition, may enable calculation of theproper roll angle to issue the attitude thruster squib command. If therehappens to be an attitude thruster at the right place at the moment,that thruster will be selected to fire. For example, in FIG. 26 ,attitude thruster number 8 is about to reach the ignition commandposition; but it has already been spent, so it cannot be selected.Thruster number 7 is available but it's too soon to be fired so it willnot be selected, either. However, as the KV 400 precesses and nutates,the ignition command angle ϕ_(ignit,cmd) and thruster number 7 may lineup and be selected to fire.

Act 12: The attitude control loop may keep track of which attitudethrusters are available to use in an availability list. When a certainattitude thruster is commanded to fire, it is removed from theavailability list.

Because of the high mass ratio of the divert thruster 610 and theattitude control thruster 1604 propellants to the KV 400 total mass, theCG location and the inertia tensor will migrate as any divert thruster610 or an attitude control thruster 1604 is firing. This is theso-called CG migration problem and it becomes more important toward theend of flight when most of the thrusters are spent.

FIG. 27 illustrates a mass and CG location update algorithm 2700. Themass and CG location update algorithm 2700 may be configured to updatethe total mass, CG location and inertia tensor as a divert thruster 610or attitude control thruster 1604 is firing. The definitions of thesymbols are:

{dot over (m)} _(i) ≡i ^(th) thruster mass burn rate<0

ir _(i) ≡i ^(th) thruster CG location

m≡total mass

r _(CG) Cg location

I _(xy) Inertia tensor element xy  (37)

The following equations are computed recursively:

$\begin{matrix}{{{\Delta m_{i}} = {{{\overset{.}{m}}_{i} \cdot {dt}} < 0}}{{m( t_{k} )} = {{m( t_{k - 1} )} + {\Delta m_{i}}}}{{r_{CG}(t)} = \frac{{{m( t_{k - 1} )}{r_{CG}( t_{k - 1} )}} + {\Delta m_{i}r_{i}}}{m( t_{k} )}}{{I_{xx}( t_{k} )} = {{I_{xx}( t_{k - 1} )} - {\sum{\Delta{m_{i}( {y_{i}^{2} + z_{i}^{2}} )}}}}}{{I_{yy}( t_{k} )} = {{I_{yy}( t_{k - 1} )} - {\sum{\Delta{m_{i}( {x_{i}^{2} + z_{i}^{2}} )}}}}}{{I_{zz}( t_{k} )} = {{I_{zz}( t_{k - 1} )} - {\sum{\Delta{m_{i}( {x_{i}^{2} + y_{i}^{2}} )}}}}}{{I_{xy}( t_{k} )} = {{I_{xy}( t_{k - 1} )} + {\sum{\Delta m_{i}x_{i}y_{i}}}}}{{I_{xz}( t_{k} )} = {{I_{xz}( t_{k - 1} )} + {\sum{\Delta m_{i}x_{i}z_{i}}}}}{{I_{yz}( t_{k} )} = {{I_{yz}( t_{k - 1} )} + {\sum{\Delta m_{i}y_{i}z_{i}}}}}} & (38)\end{matrix}$

To a large extent, this detailed description has focused on a particulartype of intercept vehicle (e.g., the eject vehicle 400). However,engagement management systems described herein may be used with manytypes of intercept vehicles 400 in which the engagement managementsystem can track the intercept vehicle 400, alter the course of theintercept vehicle 400, determine when to detonate the intercept vehicle400, or combinations thereof using commands communicated between theengagement management system and the intercept vehicle 400.

Moreover, while embodiments of the present disclosure may beparticularly suitable for use on aerial platforms, they may also be usedin other types of mobile platforms like ground-based mobile platformssuch as, for example, tanks, armored personnel carriers, personnelcarriers (e.g., Humvee and Stryker vehicles) and other mobile platformscapable of bearing embodiments of the present disclosure. Moreover,embodiments of the present disclosure may be used for relativelystationary ground-based personnel protection wherein a mobile platformmay not be involved. Accordingly, embodiments of the disclosure are notlimited to aerial applications.

While the present disclosure has been described herein with respect tocertain illustrated embodiments, those of ordinary skill in the art willrecognize and appreciate that the present invention is not so limited.Rather, many additions, deletions, and modifications to the illustratedand described embodiments may be made without departing from the scopeof the invention as hereinafter claimed along with their legalequivalents. In addition, features from one embodiment may be combinedwith features of another embodiment while still being encompassed withinthe scope of the invention as contemplated by the inventor.

1. An active protection system, comprising: an aircraft; an aerialvehicle carried by the aircraft, the aerial vehicle comprising: a rocketmotor configured to accelerate the aerial vehicle in a direction havinga component parallel to a longitudinal axis of the aerial vehicle; andalignment thrusters configured to cause the aerial vehicle to perform apitch maneuver by rotating the aerial vehicle about an axis orthogonalto the longitudinal axis of the aerial vehicle prior to the rocket motorbeing fired.
 2. The active protection system of claim 1, wherein theaerial vehicle further comprises divert thrusters configured to divertthe aerial vehicle in a direction transverse to the longitudinal axis ofthe aerial vehicle subsequent to the rocket motor being fired.
 3. Theactive protection system of claim 1, wherein the aerial vehicle furthercomprises a nose thruster module proximate a front end of the aerialvehicle.
 4. The active protection system of claim 1, wherein the aerialvehicle further comprises: a controller; and one or more sensorsoperably coupled to the controller, wherein the controller is configuredto initiate flight corrections to the aerial vehicle during flight basedon feedback from the one or more sensors.
 5. The active protectionsystem of claim 4, further comprising a guidance system carried by theaircraft and configured to transmit guidance signals to the aerialvehicle during flight of the aerial vehicle.
 6. The active protectionsystem of claim 1, further comprising a detection system carried on theaircraft and comprising at least one sensor configured to detect aerialthreats with threat range of the aircraft.
 7. The active protectionsystem of claim 6, wherein the detection system is further configured totrack aerial threats detected by the at least one sensor.
 8. The activeprotection system of claim 6, further comprising an aerial vehiclecarrier disposed on the aircraft.
 9. A method of neutralizing an aerialthreat, comprising: causing an aerial vehicle to perform a first actionof a launch sequence to leave a platform; subsequent to the first actionof the launch sequence, causing the aerial vehicle to perform a secondaction of the launch sequence, the second action comprising performing apitch maneuver by rotating the aerial vehicle about an axis orthogonalto a longitudinal axis of the aerial vehicle; and subsequent to thesecond action of the launch sequence, firing a rocket motor of theaerial vehicle to cause the aerial vehicle to accelerate in a directionhaving a component parallel to the longitudinal axis of the aerialvehicle.
 10. The method of claim 9, wherein causing an aerial vehicle toperform a first action of a launch sequence comprises releasing theaerial vehicle from the platform.
 11. The method of claim 9, whereincausing an aerial vehicle to perform a first action of a launch sequencecomprises propelling the aerial vehicle from the platform.
 12. Themethod of claim 9, wherein causing an aerial vehicle to perform a firstaction of a launch sequence comprises launching the aerial vehicle via adispenser of the platform.
 13. The method of claim 9, further comprisingperforming a divert maneuver with one or more divert thrusters of theaerial vehicle to divert the aerial vehicle one or more times subsequentto firing the rocket motor of the aerial vehicle.
 14. The method ofclaim 9, further comprising causing the aerial vehicle to perform thesecond action of the launch sequence after the aerial vehicle is aparticular distance away from the platform.
 15. The method of claim 9,wherein causing an aerial vehicle to perform a first action of a launchsequence comprises releasing the aerial vehicle from a carrier disposedon the platform into air deflected by the platform.
 16. The method ofclaim 9, further comprising initiating the first action of the launchsequence responsive to identifying the aerial threat.
 17. A method ofcountering an aerial threat, comprising: causing an aerial vehicle toexit a vicinity of an aerial platform; causing the aerial vehicle toperform a pitch maneuver by rotating the aerial vehicle about an axisorthogonal to an longitudinal axis of the aerial vehicle; and subsequentto causing the aerial vehicle to perform the pitch maneuver, firing arocket motor of the aerial vehicle to cause the aerial vehicle toaccelerate in a direction having a component parallel to thelongitudinal axis of the aerial vehicle.
 18. The method of claim 17,further comprising identifying the aerial threat as a potential threatto at least one object.
 19. The method of claim 18, wherein the at leastone object is one of the aerial platform, a ground vehicle, anotheraerial platform, or a ground structure.
 20. The method of claim 18,further comprising detonating a payload of the aerial vehicle uponinterception of the aerial threat with the aerial vehicle.